Apparatuses and methods for joining composite members and other structural members in aircraft wing boxes and other structures

ABSTRACT

Apparatuses and methods for joining composite members and other structural members in aircraft wings and other structures. An aircraft wing box structure configured in accordance with one embodiment of the invention includes a first composite member having a first surface portion positioned at an angle relative to a second surface portion of a second composite member. The wing box structure of this embodiment further includes at least one metallic joining member having an upstanding leg portion extending from a base portion. The base portion of the joining member is bonded to the first surface portion of the first composite member with a first portion of adhesive, and the upstanding leg portion of the metallic joining member is bonded to the second surface portion of the second composite member with a second portion of adhesive.

CROSS REFERENCE TO RELATED APPLICATION

This is a divisional application of U.S. patent application Ser. No.11/607,333, filed Nov. 30, 2006 now U.S. Pat. No. 7,575,194, entitledAPPARATUSES AND METHODS FOR JOINING COMPOSITE MEMBERS AND OTHERSTRUCTURAL MEMBERS IN AIRCRAFT WING BOXES AND OTHER STRUCTURES, which isherein incorporated by reference in its entirety.

TECHNICAL FIELD

The following disclosure relates generally to aircraft structures and,more particularly, to apparatuses and methods for joining compositemembers and other structural members in aircraft wings and otherstructures.

BACKGROUND

Fiber-reinforced resin materials, or “composite” materials as they arecommonly known, have relatively high strength-to-weight ratios, goodcorrosion resistance, and other beneficial properties that make themparticularly well suited for use in aerospace applications. Conventionalcomposite materials typically include glass, carbon, or polyaramidfibers in woven and non-woven configurations. In the raw material stage,the fibers can be formed into tapes, filaments, and fabric sheets thatare pre-impregnated with uncured resin. The raw materials can bemanufactured into parts by laminating them onto a mold surface, and thenapplying heat and pressure to cure the resin and harden the laminate.Composite sandwich structures can be manufactured by laminating a corematerial (e.g., a foam or honeycomb material) between two facesheetscomposed of laminated plies, tapes, and/or filaments. Facesheets canalso include one or more metallic layers.

Because of their relatively high strength-to-weight ratios, compositematerials are often used in aircraft structures to reduce weight andincrease performance. In fighter aircraft, business jets, and otherrelatively high-performance aircraft, for example, composite materialshave been used in both primary and secondary structures. In largecommercial aircraft, however, the use of composite materials hastraditionally been limited to non-critical, secondary structures, whilewing spars and other primary structures have been manufacturedpredominantly from metals such as aluminum, titanium, etc.

When used in primary structure, composite wing spars are typicallymanufactured by forming a solid laminate of fiber plies having a “C”cross-sectional shape. This relatively simple method reduces part countand lends itself well to automated lay-up procedures. One downside ofthis approach, however, is that it can be difficult to vary the plycount over the length and height of the spar. As a result, some portionsof the spar may be much thicker (and heavier) than they need to be tomeet localized structural requirements. In addition, composite sparsmanufactured in this way often have to be reinforced with stiffenerswhich are bolted or bonded to the spar web between ribs to limitbuckling. Moreover, such spars often do not meet ground plane andelectromagnetic effects (EME) requirements without the addition ofrelatively heavy ground cables to the upper and lower portions of thespar.

SUMMARY

The following summary is provided for the benefit of the reader only,and is not intended to limit the invention as set forth by the claims inany way.

The present invention is directed generally toward apparatuses andmethods for joining structural members in aircraft wing boxes and otherstructures. An aircraft structure configured in accordance with oneaspect of the invention includes first and second structural members.The first structural member is constructed of composite materials andhas a first surface portion and a second surface portion forming atapered edge portion. The tapered edge portion at least approximatelyabuts a third surface portion of the second structural member. Theaircraft structure further includes first and second metallic joiningmembers. The first metallic joining member has a first base portionpositioned adjacent to the third surface portion of the secondstructural member, and a first upstanding leg portion positionedadjacent to the first surface portion of the first structural member.The second metallic joining member has a second base portion positionedadjacent to the third surface portion of the second structural member,and a second upstanding leg portion positioned adjacent to the secondsurface portion of the first structural member. In addition to theforegoing elements, the aircraft structure additionally includes firstand second portions of adhesive. The first portion of adhesive forms afirst structural bond between the first upstanding leg portion of thefirst metallic joining member and the first surface portion of the firststructural member. The second portion of adhesive forms a secondstructural bond between the second upstanding leg portion of the secondmetallic joining member and the second surface portion of the firststructural member.

An aircraft wing structure configured in accordance with another aspectof the invention includes a plurality of wing ribs extending between afront wing spar and a rear wing spar. The front wing spar has a frontspar web, the rear wing spar has a rear spar web, and each wing rib hasa corresponding rib web constructed from composite materials. Each ribweb is fixedly attached to the front and rear spar webs by a firstmetallic joining member and a second metallic joining member. Morespecifically, the first metallic joining member has a first surfaceportion structurally bonded to the front spar web and a second surfaceportion structurally bonded to the rib web. Similarly, the secondmetallic joining member has a third surface portion bonded to the rearspar web and a fourth surface portion structurally bonded to the ribweb.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partially hidden plan view of a wing box structureconfigured in accordance with an embodiment of the invention.

FIG. 2 is an enlarged cross-sectional view of the wing box structure ofFIG. 1, taken substantially along line 2-2 in FIG. 1.

FIG. 3 is an enlarged cross-sectional view of a portion of the wing boxstructure illustrated in FIG. 2.

FIG. 4 is an enlarged cross-sectional view of the wing box structure ofFIG. 1, taken substantially along line 4-4 in FIG. 1.

FIG. 5 is an enlarged cross-sectional view of the wing box structure ofFIG. 1, taken substantially along line 5-5 in FIG. 1.

FIG. 6 is an enlarged cross-sectional view of a portion of the wing boxstructure illustrated in FIG. 5.

FIG. 7 is an enlarged cross-sectional view of a portion of the wing boxstructure illustrated in FIG. 4.

FIG. 8 is an enlarged cross-sectional view of another portion of thewing box structure illustrated in FIG. 4.

FIG. 9 is an enlarged cross-sectional view of a wing box structure jointconfigured in accordance with another embodiment of the invention.

DETAILED DESCRIPTION

The following disclosure describes apparatuses and methods for joiningcomposite members and other structural members in aircraft wing boxesand other structures. Certain details are set forth in the followingdescription and in FIGS. 1-9 to provide a thorough understanding ofvarious embodiments of the invention. Other details describingwell-known methods, structures and systems often associated with themanufacture and assembly of composite parts and aircraft structures arenot set forth in the following disclosure to avoid unnecessarilyobscuring the description of the various embodiments of the invention.

Many of the details, dimensions, angles and other features shown in theFigures are merely illustrative of particular embodiments of theinvention. Accordingly, other embodiments can have other details,dimensions, angles and features without departing from the spirit orscope of the present invention. In addition, further embodiments of theinvention can be practiced without several of the details describedbelow.

In the Figures, identical reference numbers identify identical, or atleast generally similar, elements. To facilitate the discussion of anyparticular element, the most significant digit or digits of anyreference number refer to the Figure in which that element is firstintroduced. For example, element 110 is first introduced and discussedwith reference to FIG. 1.

FIG. 1 is a partially hidden top view of an aircraft wing 102 having awing box 100 configured in accordance with an embodiment of theinvention. The wing box 100 includes a series of wing ribs 108(identified individually as ribs 108 a-j) extending between a front wingspar 106 and a rear wing spar 104. An upper skin panel 110 and a lowerskin panel 112 are attached to the front spar 106, the rear spar 104,and the wing ribs 108 to form an enclosed box structure.

The wing box 100 is the primary load carrying structure of the wing 102.In this regard, the wing box 100 can carry a number of importantaircraft systems including, for example, fuel tanks, engine supports,control surface actuation systems, landing gears, etc. As described ingreater detail below, in this embodiment the front spar 106, the rearspar 104, the upper skin panel 110, and the lower skin panel 112 can beindividually manufactured from composite materials and joined togetherwith metallic joining members that are bonded to the individualcomposite parts to form the wing box 100. In other embodiments, however,the front spar 106, the rear spar 104, the upper skin panel 110, and/orthe lower skin panel 112 can be manufactured, or at least partiallymanufactured, from metals, such as aluminum, titanium, and/or steel. Forexample, in various embodiments of the invention, the front spar 106and/or the rear spar 104 can be manufactured from metal usingconventional techniques known in the art. In yet other embodiments, oneor more of the structural members identified above can be manufacturedfrom composite materials but can also include one or more metallicfacesheets or other elements. For example, in various embodiments thefront spar 106 and/or the rear spar 104 can include composite sandwichstructures with one or more metallic facesheets for additional strengthand/or other reasons.

FIG. 2 is an enlarged cross-sectional view of the wing box 100 takensubstantially along line 2-2 in FIG. 1. In one aspect of thisembodiment, the upper skin panel 110 can include a core 218 sandwichedbetween a first facesheet 214 a and a second facesheet 214 b. Thefacesheets 214 can be constructed from lay-ups of fiber-reinforced resinmaterials. Such materials can include, for example, graphite-reinforcedepoxy materials in fabric, tape, tow, filament and/or other suitableforms, as well as other suitable fiber-reinforced resin materials. Inother embodiments, the facesheets can include metallic materials, suchas aluminum, titanium, and/or steel in skin, panel, and/or other forms.The core 218 can include various types of honeycomb materials, such asNOMEX® aramid fiber honeycomb, aluminum honeycomb, and graphite/epoxy,as well as various types of open or closed cell foam and/or othersuitable core materials.

In one embodiment, the upper skin panel 110 can be constructed by usingan automated lay-up process to arrange a plurality of fabric and/or tapeplies against a tool surface (not shown) to form the first facesheet 214a. A first layer of adhesive can then be applied to the first facesheet214 a, and the core 218 can be positioned on the first layer ofadhesive. A second layer of adhesive can then be applied to the surfaceof the core 218, and a similar automated lay-up process can be used tooverlay the core 218 with additional fabric and/or tape plies to formthe second facesheet 214 b. The thickness of the core 218 can taper downtoward the front spar 106 and the rear spar 104 so that the face sheets214 can form a solid laminate in these areas. This assembly can then bevacuum-bagged and positioned in an evacuated and/or elevated-temperatureenvironment (e.g., an autoclave) for curing. In other embodiments, theupper skin panel 110 can be manufactured by other compositemanufacturing methods known in the art.

In another aspect of this embodiment, the upper skin panel 110 caninclude one or more conduits 220 through which electrical wiring,hydraulic lines and/or other wing systems can pass. The lower skin panel112 can be at least generally similar in structure as the upper skinpanel 110. In the illustrated embodiment, however, the lower skin panel112 can include a removable panel 226 to provide access to the interiorof the wing box 100 for maintenance, inspection, etc.

The front spar 106 includes a front spar web 228 that extends between alower edge portion 250 and an upper edge portion 252. The front spar web228 can be at least generally similar in construction as the upper skinpanel 110 described above. In this regard, the front spar web 228 caninclude a core 234 sandwiched between a first facesheet 230 a and asecond facesheet 230 b. In this embodiment, the face sheets 230 arebonded, laminated, or otherwise joined together along the upper andlower edges of the front spar web 228 so that both the lower edgeportion 250 and the upper edge portion 252 are tapered as shown in FIG.2. The facesheets 230 can include various types of composite materials,such as graphite/epoxy fabric and tape materials. In addition, thefacesheets 230 can also include various types of metallic materials. Themetallic materials can be used in conjunction with the compositematerials (e.g., bonded to the composite facesheets) or used in place ofcomposite facesheets. The core 234 can include suitable types ofhoneycomb, foam, and other known materials. In one aspect of thisembodiment, the front spar web 228 can be manufactured using anautomated, flat lay-up process. This process can reduce manufacturingcosts because the first facesheet 230 a is laid-up against a relativelyflat tool surface, rather than a curved tool surface. The rear spar 104includes a rear spar web 236 extending between a lower edge portion 238and an upper edge portion 240. The rear spar 104 can be at leastgenerally similar in construction as the front spar 106.

Although the front spar 106 and the rear spar 104 can include compositesandwich structures as described above, the present invention is notlimited to this particular embodiment. Indeed, in other embodiments, thefront spar 106 and/or the rear spar 104 can be machined, built-up,and/or otherwise fabricated from metallic materials using conventionalspar manufacturing methods known in the art. In such embodiments, thefront spar web 228 and the rear spar web 236 could be formed from solidmetal sheets, panels, and/or other forms.

The wing box 100 can also include a plurality of metallic joiningmembers that are adhesively bonded to the front spar 106, the rear spar104, the upper skin panel 110 and the lower skin panel 112 tostructurally attach the composite members together. For example, thelower edge portion 238 of the rear spar web 236 can be fixedly attachedto the lower skin panel 112 with a first spar joining member 241 and asecond spar joining member 242. Similarly, the upper edged portion 240of the rear spar web 238 can be fixedly attached to the upper skin panel110 by a third spar joining member 243 and a fourth spar joining member244. Moving to the front spar 106, the lower edge portion 250 of thefront spar web 228 can be fixedly attached to the lower skin panel 112by a fifth spar joining member 245 and a sixth spar joining member 246.Similarly, the upper edge portion 252 of the front spar web 228 can befixedly attached to the upper skin panel 110 by a seventh spar joiningmember 247 and an eighth spar joining member 248. Each of thesestructural joints is described in more detail below.

FIG. 3 is an enlarged cross-sectional view illustrating the joint wherethe rear spar web 236 at least approximately abuts the lower skin panel112. The first joining member 241 includes a first upstanding legportion 346 and a first base portion 348. The second joining member 242includes a second upstanding leg portion 350 and a second base portion352. In this embodiment, the lower edge portion 240 of the rear spar web236 is sandwiched between the first upstanding leg portion 346 and thesecond upstanding leg portion 350, and the first base portion 348 issandwiched between the second base portion 352 and the lower skin panel112. Adhesive 344 can be applied to the mating surfaces of the rear sparweb 236, the first joining member 241, the second joining member 242,and the lower skin panel 112 to bond the parts together. At least thedistal end portions of the first upstanding leg portion 346 and/or thesecond upstanding leg portion 350 can be tapered as shown in FIG. 3 forefficient load transfer between the joining members and the rear sparweb 236. Tapering in this manner can also reduce peak loads at thedistal ends of the upstanding leg portions. In some embodiments, one ormore fasteners 354 can be installed through the first base portion 348,the second base portion 352, and the lower skin panel 112 (in additionto, or in place of, the adhesive 344) to structurally attach the firstjoining member 241 and the second joining member 242 to the lower skinpanel 112. Various types of fasteners can be used for this purposeincluding, for example, bolts, lock-pins, rivets, etc.

The joining members 241 and 242 can be made from various types ofmetallic materials including, for example, aluminum, titanium, stainlesssteel, etc. In one embodiment, the joining members 241 and 242 can bebonded to the rear spar web 236 and/or the lower skin panel 112 with“bond on demand” technology. With “bond on demand” technology, theadhesive is applied to the bonding surfaces, but the adhesive does notcure or harden until it is activated within the bond line with radiationenergy (e.g., X-Ray, electron beam, Ultraviolet and/or other lightenergy, etc.). This technique allows the parts to be adjusted duringfinal assembly, and it avoids putting detrimental heat into the partsduring the adhesive cure cycle. This method also allows bonding of largeassemblies that might otherwise not be able to fit into an autoclave oroven for curing. Types of adhesives that can be used for this purposeinclude acrylate or epoxy adhesives, such as AP299 adhesive, etc. Inother embodiments, other adhesives (e.g., other thermally curedadhesives) can be used to bond the rear spar web 236 to the lower skinpanel 112. Such adhesives can include, for example, epoxy adhesives(e.g., HYSOL® EA9696 epoxy adhesive, HYSOL® EA9380 epoxy adhesive, etc.)and other suitable adhesives known in the art for structurally joiningmetallic materials to composite materials. In some embodiments, thebonding methods and/or systems disclosed in U.S. patent application Ser.No. 11/154,522, filed Jun. 6, 2005, and entitled “CompositeReinforcement of Metallic Structural Elements,” can be used to bond twoor more of the structural parts described herein together. U.S. patentapplication Ser. No. 11/154,522 is incorporated herein in its entiretyby reference.

FIG. 4 is an enlarged cross-sectional view of the wing box 100 takensubstantially along line 4-4 in FIG. 1, and FIG. 5 is an enlargedcross-sectional view of the wing box 100 taken substantially along line5-5 in FIG. 1. These Figures illustrate a portion of the wing box 100that includes the wing rib 108 d. Referring to FIGS. 4 and 5 together,the wing rib 108 d can include a rib web 460 that is generally similarin construction as the front spar web 228 described above with referenceto FIG. 2. More specifically, the rib web 460 can include a core 534sandwiched between a first facesheet 530 a and a second facesheet 530 b.The facesheets 530 can include various types of composite materials,such as graphite/epoxy fabric and tape materials. The core 534 caninclude suitable types of honeycomb, foam, and other known materials. Inone aspect of this embodiment, the rib web 460 can be manufactured usingan automated, flat lay-up process. In other embodiments, the rib web 460can have other structural configurations and can be manufactured withother suitable methods known in the art. For example, in one otherembodiment, the rib web 460 can be a solid laminate of compositematerial.

The rib web 460 can carry a conduit 420 for passage of various wingsystems (e.g. electrical systems, hydraulics systems, etc.). The rib web460 can also include one or more holes 422 for passage of air (or fuelif the wing rib 108 is positioned in a wing fuel tank). The rib web 460includes an upper edge portion 462, a lower edge portion 464, a forwardedge portion 466, and an aft edge portion 468. As described in greaterdetail below with reference to FIGS. 5-8, the lower edge portion 464 canbe attached to the lower skin panel 112 with a first rib joining member451 and a second rib joining member 452. Similarly, the upper edgeportion 462 can be attached to the upper skin panel 110 with a third ribjoining member 453 and a fourth rib joining member 454. In addition, therear edge portion 468 can be attached to the rear spar web 104 with afifth rib joining member 455, and the forward edge portion 466 can beattached to the front spar web 106 with a sixth rib joining member 456.

FIG. 6 is taken from FIG. 5, and is an enlarged cross-sectional viewillustrating the joint between the rib web 460 and the lower skin panel112 in more detail. As this view illustrates, the first rib joiningmember 451 can include a first upstanding leg portion 646 and a firstbase portion 648. The second rib joining member 452 can include a secondupstanding leg portion 650 and a second base portion 652. In thisembodiment, the lower edge portion 464 of the rib web 460 is sandwichedbetween the first upstanding leg portion 646 and the second upstandingleg portion 650, and the first base portion 648 is sandwiched betweenthe second base portion 652 and the lower skin panel 112. The adhesive344 can be applied to the mating surfaces of the rib web 460, the firstupstanding leg portion 646, and the second upstanding leg portion 650 tobond the rib web 460 to the joining members (using, e.g., bond on demandtechnology as described above, or another suitable method known in theart). The adhesive 344 can also be used to bond the first base portion648 to the second base portion 652 and to an inner facesheet 616 of thelower skin panel 112. In addition or alternatively, one or morefasteners 654 can also be used to structurally attach the first ribjoining member 451 and the second rib joining member 452 to the lowerskin panel 112. As FIG. 6 illustrates, the inner facesheet 616 of thelower skin panel 112 can include additional composite material in abuilt-up area 602 adjacent to the rib web 460 to efficiently transferloads from the rib web 460 into the lower skin panel 112.

FIG. 7 is an enlarged cross-sectional view taken along line 7-7 in FIG.4, and illustrates the joint where the wing rib 108 d at leastapproximately abuts the rear spar 104. The fifth rib joining member 455can include an upstanding leg portion 741 and a base portion 742. Theupstanding leg portion 741 can be bonded to the rib web 460 with theadhesive 344. Similarly, the base portion 742 can be bonded to the rearspar web 236 with the adhesive 344. In addition or alternatively, one ormore fasteners 754 can be used to structurally attach the fifth joiningmember 455 to the rear spar web 236.

In the embodiment illustrated in FIG. 7, only a single joining member,i.e., the fifth rib joining member 455, is used to attach the rib web460 to the rear spar web 236. In other embodiments, however, two or morejoining members can be used for this purpose. For example, in anotherembodiment, two joining members in the configuration illustrated in FIG.3 can be used to attach the rib web 460 to the rear spar web 236.

FIG. 8 is taken from FIG. 4, and is an enlarged cross-sectional viewillustrating the joint between the wing rib 108 d, the rear spar 104,and the lower skin panel 112 in more detail. As this view illustrates,one or more structural fasteners 854 (e.g., bolts, lock pins, rivets,etc.) can be used to attach the end portion of the first rib joiningmember 451 to the end portion of the second rib joining member 452and/or the adjacent end portion of the fifth rib joining member 455.Such fasteners may be required or advantageous in other areas where therib joining members transfer relatively high loads, as may be the casewhere a fitting 870 (e.g., a control surface hinge or gear fitting) isattached to the backside of the rear spar 104 adjacent to the wing rib108 d.

FIG. 9 is an enlarged cross-sectional view illustrating a structuraljoint (e.g., a composite rib/spar joint) configured in accordance withanother embodiment of the invention. Here, a first composite member 960(e.g., a composite rib web, spar web, etc.) is bonded to a secondcomposite member 910 (e.g., a wing skin panel, fuselage skin panel,control surface panel, etc.) by a first joining member 941 a and asecond joining member 941 b. Each of the joining members 941 includes anupstanding leg portion 946 (identified individually as a firstupstanding leg portion 946 a and a second upstanding leg portion 946 b)which extends from a corresponding base portion 948 (identifiedindividually as a first base portion 948 a and a second base portion 948b). The first and second joining members 941 are positioned back-to-backso that the corresponding base portions 948 extend outwardly to form a“T.” The first composite member 960 can be bonded to the upstanding legportions 946 with the adhesive 344. The base portions 948 can also bebonded to the second composite member 910 with the adhesive 344. Inaddition or alternatively, a plurality of fasteners 954 can be used tostructurally attach the joining members 941 to the second compositemember 910.

Although various joining member configurations have been described abovefor purposes of illustration, other joining member configurations can beused to bond wing box members and other structures together as disclosedherein. Such configurations can include, for example, one-piece joiningmembers which have a groove to receive an edge portion of at least onecomposite member. These one-piece joining members can have “L” shapesresembling the combined two-piece arrangement of FIG. 3, or a “T” shapesresembling the two-piece configuration of FIG. 9. In other embodiments,still further joining member configurations can be used to assemblecomposite wing boxes and other structures without departing from thepresent disclosure.

There are a number of advantages associated with various embodiments ofthe invention described above. One advantage of bonding composite ribs,spars, and/or skin panels together is that it spreads the load over alarger area, providing a uniform load distribution across the jointwithout the peak loads associated with bolted joints. This method alsoseals the joint eliminating or reducing leak paths. Furthermore, the useof dissimilar materials (i.e., composites and metals) allowsoptimization of structural functions. For example, metal flanges can besized for out-of-plane loads, while composite webs can be sized forshear transfer. In addition, the “flat” composite webs described hereincan be manufactured by automated lay-up processes, which can improvequality and reduce manufacturing costs as compared to spar and webmembers which are hand-formed into solid laminates with “C”cross-sectional shapes.

From the foregoing, it will be appreciated that specific embodiments ofthe invention have been described herein for purposes of illustration,but that various modifications may be made without deviating from thespirit and scope of the various embodiments of the invention. Forexample, although aspects of the invention have been described above inthe context of an aircraft wing, in other embodiments, the structuralmethods and apparatuses described above can be used in the constructionof other types of composite members (e.g., fuselage members, empennagemembers, etc.) in aircraft and other structures. Further, while variousadvantages associated with certain embodiments of the invention havebeen described above in the context of those embodiments, otherembodiments may also exhibit such advantages, and not all embodimentsneed necessarily exhibit such advantages to fall within the scope of theinvention. Accordingly, the invention is not limited, except as by theappended claims.

1. An aircraft wing box structure comprising: a front wing spar having afront spar web, wherein the front spar web is constructed from compositematerials; a rear wing spar having a rear spar web, wherein the rearspar web is constructed from composite materials; a plurality of wingribs extending between the front wing spar and the rear wing spar,wherein each wing rib has a corresponding rib web constructed fromcomposite materials, wherein each rib web has a front portion and a rearportion, and wherein each rib web is fixedly attached to the front sparweb by: a first metallic joining member having a first base portion anda first upstanding leg portion, wherein the first base portion includesa first surface portion structurally bonded to the front spar web, andwherein the first upstanding leg portion includes a second surfaceportion structurally bonded to the front portion of the rib web; and asecond metallic joining member having a second base portion and a secondupstanding leg portion, wherein the second upstanding leg portionincludes a third surface portion structurally bonded to the frontportion of the rib web opposite the first upstanding leg of the firstmetallic joining member, and wherein the first base portion of the firstmetallic joining member is sandwiched between the second base portion ofthe second metallic joining member and the front spar web.
 2. Theaircraft wing box structure of claim 1 wherein the front and rear sparwebs are generally flat, and wherein each of the front and rear sparwebs has a tapered upper edge portion extending in a first direction anda tapered lower edge portion extending in a second direction opposite tothe first direction.
 3. The aircraft wing box structure of claim 1wherein one side of each rib web is at least generally flat and theother side of each rib web is convex.
 4. The aircraft wing box structureof claim 1 wherein each rib web includes a core portion sandwichedbetween a first face sheet and a second face sheet.
 5. The aircraft wingbox structure of claim 1 wherein the front spar web includes a firstcore portion sandwiched between a first face sheet and a second facesheet, and the rear spar web includes a second core portion sandwichedbetween a third face sheet and a fourth face sheet.
 6. The aircraft wingbox structure of claim 1 wherein each rib web is further fixedlyattached to the rear spar web by: a third metallic joining member havinga fourth surface portion structurally bonded to the rear spar web and afifth surface portion structurally bonded to the rear portion of the ribweb.
 7. The aircraft wing box structure of claim 1 wherein each rib webfurther includes a fourth surface portion opposite a fifth surfaceportion, wherein the second surface portion is structurally bonded tothe fourth surface portion of the rib web, and wherein the third surfaceportion is structurally bonded to the fifth surface portion of the ribweb.
 8. An aircraft wing structure comprising: a first composite wingspar; a second composite wing spar; a composite rib extending at leastpartially between the first and second composite wing spars, thecomposite rib having a front edge portion positioned proximate to thefirst composite wing spar and a rear edge portion positioned proximateto the second composite wing spar, wherein the front edge portionincludes a first side facing opposite a second side; first means forfixedly attaching the first side of the front edge portion of thecomposite rib to the first composite wing spar in the absence of anystructural fasteners extending through the first composite wing spar,wherein the first means includes a first base portion bonded to thefirst composite wing spar; and second means for fixedly attaching thesecond side of the front edge portion of the composite rib to the firstcomposite wing spar in the absence of any structural fasteners extendingthrough the first composite wing spar, wherein the second means includesa second base portion, and wherein the first base portion is sandwichedbetween the second base portion and the first composite wing spar. 9.The aircraft wing structure of claim 8 wherein the front edge portion ofthe composite wing rib is tapered, and wherein the first and secondmeans include means for receiving the tapered front edge portion. 10.The aircraft wing structure of claim 8 wherein the composite rib furtherincludes an upper edge portion and a lower edge portion, and wherein theaircraft wing structure further comprises: a first composite skin panelcovering at least a portion of the first composite wing spar and thesecond composite wing spar proximate to the upper edge portion of thecomposite rib; a second composite skin panel covering at least a portionof the first composite wing spar and the second composite wing sparproximate to the lower edge portion of the composite rib; means forfixedly attaching the upper edge portion of the composite rib to thefirst composite skin panel in the absence of any structural fastenersextending through the composite rib; and means for means for fixedlyattaching the lower edge portion of the composite rib to the secondcomposite skin panel in the absence of any structural fastenersextending through the composite rib.
 11. The aircraft wing structure ofclaim 8 wherein the composite rib further includes a core portionsandwiched between a first face sheet and a second face sheet, whereinthe first and second face sheets are bonded together along the frontedge portion of the composite wing rib to form a tapered front edgeportion, wherein the first and second face sheets are bonded togetheralong the rear edge portion of the composite wing rib to form a taperedrear edge portion, wherein the first and second face sheets are bondedtogether along an upper edge portion of the composite wing rib to form atapered upper edge portion, wherein the first and second face sheets arebonded together along a lower edge portion of the composite wing rib toform a tapered lower edge portion, and wherein the aircraft wingstructure further comprises: a first composite skin panel covering atleast a portion of the first composite wing spar and the secondcomposite wing spar proximate to the tapered upper edge portion of thecomposite rib; a second composite skin panel covering at least a portionof the first composite wing spar and the second composite wing sparproximate to the tapered lower edge portion of the composite rib; meansfor fixedly attaching the tapered upper edge portion of the compositerib to the first composite skin panel in the absence of any structuralfasteners extending through the composite rib; and means for means forfixedly attaching the tapered lower edge portion of the composite rib tothe second composite skin panel in the absence of any structuralfasteners extending through the composite rib.
 12. An aircraft wing boxstructure comprising: a wing spar having a spar web; a wing rib having arib web formed by a core portion sandwiched between a first facesheetand a second facesheet, wherein the first and second facesheets arejoined together along an outer boundary to form a tapered edge portionthat at least approximately abuts the spar web; a first metallic joiningmember having a first base portion positioned adjacent to the spar weband a first upstanding leg portion positioned adjacent to the firstfacesheet of the rib web; a second metallic joining member having asecond base portion positioned adjacent to the spar web and a secondupstanding leg portion positioned adjacent to the second facesheet ofthe rib web, wherein the outer boundary of the tapered edge portion issandwiched between the first upstanding leg portion and the secondupstanding leg portion; a first portion of adhesive forming a firststructural bond between the first upstanding leg portion of the firstmetallic joining member and the first facesheet of the rib web; and asecond portion of adhesive forming a second structural bond between thesecond upstanding leg portion of the second metallic joining member andthe second facesheet of the rib web, wherein the first base portion ofthe first metallic joining member is sandwiched between the second baseportion of the second metallic joining member and the spar web.
 13. Theaircraft wing box structure of claim 12 wherein the spar web isconstructed from composite materials.
 14. The aircraft wing boxstructure of claim 12, further comprising: a third portion of adhesiveforming a third structural bond between the first base portion of thefirst metallic joining member and the spar web; and a fourth portion ofadhesive forming a fourth structural bond between the second baseportion of the second metallic joining member and the first base portionof the first metallic joining member.
 15. The aircraft wing boxstructure of claim 12 wherein the tapered edge portion of the rib web isbonded to the first upstanding leg portion of the first metallic joiningmember and the second upstanding leg portion of the second metallicjoining member in the absence of any structural fasteners extendingthrough the first and second upstanding leg portions and the taperededge portion.
 16. The aircraft wing box structure of claim 12 whereinthe spar web includes a second core portion sandwiched between a thirdfacesheet and a fourth facesheet.
 17. The aircraft wing box structure ofclaim 12 wherein the wing spar is a first wing spar having a first sparweb, wherein the tapered edge portion of the rib web is a first taperededge portion and the rib web further includes a second tapered edgeportion opposite the first tapered edge portion, and wherein theaircraft wing box structure further comprises: a second wing spar havinga second spar web, wherein the second tapered edge portion of the wingrib at least approximately abuts the second spar web; a third metallicjoining member having a third base portion positioned adjacent to thesecond spar web and a third upstanding leg portion positioned adjacentto the first facesheet of the rib web; a fourth metallic joining memberhaving a fourth base portion positioned adjacent to the second spar weband a fourth upstanding leg portion positioned adjacent to the secondfacesheet of the rib web, wherein the second tapered edge portion issandwiched between the third upstanding leg portion and the fourthupstanding leg portion; a third portion of adhesive forming a thirdstructural bond between the third upstanding leg portion of the thirdmetallic joining member and the first facesheet of the rib web; and afourth portion of adhesive forming a fourth structural bond between thefourth upstanding leg portion of the fourth metallic joining member andthe second facesheet of the rib web.
 18. The aircraft wing box structureof claim 17 wherein the wing rib further includes an upper edge portionand a lower edge portion, and wherein the aircraft wing box structurefurther comprises: a first composite skin panel covering at least aportion of the first wing spar and the second wing spar proximate theupper edge portion of the wing rib; a second composite skin panelcovering at least a portion of the first wing spar and the second wingspar proximate the lower edge portion of the wing rib; means for fixedlyattaching the upper edge portion of the wing rib to the first compositeskin panel in the absence of any structural fasteners extending throughthe wing rib; and means for means for fixedly attaching the lower edgeportion of the wing rib to the second composite skin panel in theabsence of any structural fasteners extending through the wing rib.